This invention relates generally to film cooled combustor liners for use in gas turbine engines and more particularly to such combustor liners having regions with closely spaced cooling holes.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. The fuel is injected into the combustor through fuel tubes located at uniformly spaced injection points around the combustor. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. Combustors used in aircraft engines typically include inner and outer combustor liners to protect the combustor and surrounding engine components from the intense heat generated by the combustion process. A variety of approaches have been proposed to cool combustor liners so as to allow the liners to withstand greater combustion temperatures. One such approach is multi-hole film cooling wherein a thin layer of cooling air is provided along the combustion side of the liners by an array of very small cooling holes formed through the liners. Multi-hole film cooling reduces the overall thermal load on the liners because the mass flow through the cooling holes dilutes the hot combustion gas next to the liner surfaces, and the flow through the holes provides convective cooling of the liner walls.
In the assembled combustor, certain portions of the combustor liners are aligned with the injection points defined by the circumferential location of the center of the fuel tubes. These locations are hereinafter referred to as xe2x80x9ccup centersxe2x80x9d. In operation, the flow of combustion gases past these circumferential locations create xe2x80x9chot streaksxe2x80x9d of locally increased material temperatures. The portions of the combustor liners subject-to hot streaks can exhibit oxidation, corrosion and low cycle fatigue (LCF) failures after return from field use.
Accordingly, there is a need for a combustor liner in which cooling film effectiveness is increased in the areas of the liner that are subject to unusually high temperatures and resulting material distress.
The above-mentioned need is met by the present invention, which provides a gas turbine combustor liner made up of a shell having cooling holes formed therein, a group of which are disposed upstream of the dilution holes and divided into two sub-groups. The second sub-group of this group of cooling holes is located in circumferential alignment with a hot streak and are more closely spaced than the cooling holes of the first sub-group. The shell may also have additional cooling hole groups disposed between dilution holes in the liner. The additional groups are arranged so as to provide a converging flow in the circumferential direction to provide enhanced cooling to the area of the liner downstream of the dilution holes.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.